1. Field of Endeavor
The present invention relates to airfoil blades or vanes for gas turbine engines and more particularly to an improved cooling of such blades or vanes.
2. Brief Description of the Related Art
In the field of gas turbine engines, hot combustion gases flow from a combustion chamber through one or more turbines. The hot gases provide power for one or more compressors and output power which can be used for other purposes. The turbine blades and vanes therefore have to withstand the high temperatures of the hot gases without losing operating efficiency. This can be achieved by cooling.
Various methods of internal cooling of turbine blades to keep blade temperatures within certain limits are known in the state of the art. Known methods of cooling blades and vanes are performed by supplying passages within the blades or vanes with pressurized cooling air derived from the compressor. Cooling techniques include a so-called “serpentine cooling” circuit of series-connected, longitudinally oriented passages producing serpentine flow which increase cooling effectiveness by extension of the length of the coolant flow path. Serpentine cooling is efficient by reusing the cooling air in successive longitudinal passages of the circuit. By providing openings in the webs separating the passages, cooling air streaming through the inner passages can provide convective and/or “impingement cooling” of the blade or vane before being exhausted. Impingement cooling has a high heat transfer, but can be wasteful, as the cooling air is not reused. A further known cooling method, useful for the cooling of the external wall surfaces of the airfoil, so-called “film-cooling”, is achieved by providing holes in the airfoil surface of the blade or vane. The combination of impingement cooling and film cooling is sometimes difficult as the high pressure of the cooling air necessary for impingement cooling can lead to separation of the film cooling flow from the external airfoil surface, thus reducing the cooling efficiency.
Various designs of airfoil portions are known in the state of the art. For example, the so-called “double wall concept” is known, in which at least one longitudinal wall extending parallel to and between the suction side wall and the pressure side wall of an airfoil portion. On the other hand, the mere connection of the suction side wall and the pressure side wall by longitudinal webs is known as the so-called “girder concept”. Most advanced blading technologies striving to allow higher firing temperatures allow both impingement cooling and film cooling.
U.S. Pat. No. 5,246,340 discloses an airfoil portion manufactured in two casting halves having radial internal ribs or webs with cooling air cross-over holes, said webs dividing the interior of the blade or vane in to a plurality of cavities. U.S. Patent Application Publication Nos. 2003/0133797 and 2007/0172355 disclose airfoil portions with webs connecting the suction side wall and the pressure side wall, wherein the webs are arranged angled to each other. A combination of the above mentioned double wall concept and the girder concept is seen, e.g., in U.S. Pat. No. 5,660,524. Here, the inner walls include walls which extend between and are monolithic with a portion of the outer walls. Also, a central insert is provided, forming two impingement chambers between the central insert and the external walls.